XFOIL Version 6.94 Calculated polar for: lb3000 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2655 0.04944 0.03863 -0.0351 1.0001 0.2738 -2.750 -0.2385 0.04828 0.03713 -0.0362 1.0001 0.2767 -2.500 -0.2117 0.04735 0.03588 -0.0372 1.0001 0.2811 -2.250 -0.1879 0.04652 0.03502 -0.0374 1.0001 0.2864 -2.000 -0.1626 0.04593 0.03429 -0.0378 1.0001 0.2913 -1.750 -0.1368 0.04552 0.03364 -0.0382 1.0001 0.2968 -1.500 -0.1103 0.04526 0.03314 -0.0386 1.0001 0.3040 -1.250 -0.0857 0.04511 0.03302 -0.0388 1.0001 0.3165 -1.000 -0.0603 0.04506 0.03297 -0.0391 1.0001 0.3320 -0.750 -0.0345 0.04509 0.03303 -0.0395 1.0001 0.3511 -0.500 -0.0085 0.04510 0.03326 -0.0398 1.0001 0.3783 -0.250 0.0189 0.04492 0.03370 -0.0405 1.0001 0.4414 0.000 0.0182 0.04218 0.03353 -0.0335 1.0001 0.9963 0.250 0.0456 0.04338 0.03356 -0.0345 1.0001 0.9999 0.500 0.0654 0.04465 0.03450 -0.0348 1.0001 0.9999 0.750 0.1218 0.04579 0.03527 -0.0420 0.9835 0.9999 1.000 0.2230 0.04654 0.03561 -0.0563 0.9461 0.9999 1.250 0.3122 0.04649 0.03532 -0.0673 0.9068 0.9999 1.500 0.4010 0.04556 0.03428 -0.0768 0.8634 0.9999 1.750 0.4823 0.04413 0.03278 -0.0832 0.8150 0.9999 2.000 0.5449 0.04265 0.03120 -0.0848 0.7658 0.9999 2.250 0.5811 0.04193 0.03037 -0.0825 0.7163 0.9999 2.500 0.6092 0.04145 0.02975 -0.0799 0.6729 0.9999 2.750 0.6385 0.04080 0.02893 -0.0777 0.6399 0.9999 3.000 0.6677 0.04020 0.02818 -0.0759 0.6126 0.9999 3.250 0.6980 0.03970 0.02748 -0.0744 0.5908 0.9999 3.500 0.7267 0.03982 0.02742 -0.0734 0.5716 0.9999 3.750 0.7543 0.04030 0.02777 -0.0728 0.5541 0.9999 4.000 0.7821 0.04088 0.02820 -0.0723 0.5388 0.9999 4.250 0.8142 0.04074 0.02776 -0.0713 0.5265 0.9999 4.500 0.8366 0.04262 0.02969 -0.0717 0.5123 0.9999 4.750 0.8640 0.04386 0.03084 -0.0717 0.5027 0.9999 5.000 0.8851 0.04618 0.03324 -0.0722 0.4928 0.9999